The invention relates to an aircraft with a fuselage designed essentially as an aerostatic lift body and with combined lift and propulsion devices which are articulated on the fuselage, are provided with propellers and form propulsion units and which in each case are tiltable between a lift position, in which the respective propeller rotation plane is essentially horizontal and the output shaft of the associated drive, said output shaft acting on the respective propeller shaft, is essentially vertical, and a propulsion position, in which the respective propeller rotation plane is essentially vertical and the output shaft of the associated drive, said output shaft acting on the respective propeller shaft, is essentially horizontal.
Aerostatic aircraft provided with translational propulsion systems have long been known, for example, as Zeppelins. These airships can take off and land essentially vertically by virtue of their aerostatic lift behavior, but, since they are predominantly lighter than air, they have to be restrained whenever they are held on the ground. Moreover, these aircraft are relatively sluggish to control, since, because of their slow speed and their small aerodynamic control surfaces, they have low control authority, that is to say high reaction inertia to control movements. Admittedly, airships have recently become known, which have main propulsion systems tiltable comparatively slowly about a transverse axis and laterally acting auxiliary propulsion systems for assisting the aerodynamic control, said propulsion systems reducing the turning circle of an airship, but it has not been possible for the airship to achieve xe2x80x9cPinpointxe2x80x9d maneuverability. Another disadvantage of airships with an elongated streamlined body shape is their extremely high sensitivity to crosswind and the associated requirement to restrain the airship on the ground in such a way that it can turn into the wind in the same way as a boat at a buoy, anchor masts generally being necessary for this purpose.
Furthermore, vertical take-off aircraft are known, the engines of which are tiltable out of a vertical lift position with a horizontal propeller rotation plane and a horizontal propulsion position with a vertical propeller rotation plane. One problem of these vertical take-off aircraft with tiltable engines is the control of the gyroscopic forces which occur when the engines are tilted and which have to be supported via solid supporting structures on the aircraft wings and on the fuselage. On account of these gyroscopic forces, the tilting of the engines can take place only relatively slowly. These vertical take-off aircraft are likewise controlled essentially via aerodynamic control devices. Since, during the vertical take-off of these aircraft, the engines alone have to generate the lift of the aircraft as a whole, the load capable of being transported in addition to the aircraft""s own weight is very limited.
The object of the present invention, therefore, is to provide a generic aircraft which combines the advantages of an aerostatic aircraft and the advantages of a vertical take-off aircraft and which is therefore capable of transporting large loads even over relatively long ranges, and which, at the same time, can land quickly and accurately, without a special infrastructure on the ground being required for this purpose.
This object is achieved, according to the defining part of claim 1, in that the respective propeller rotation plane has all-round inclinability relative to the output shaft of the associated drive, said output shaft acting on the propeller shaft.
This inclinability of the propeller rotation plane, in addition to the fundamentally provided tiltability of the propulsion system about a transverse axis, allows thrust vector control of the aircraft, said thrust vector control reacting quickly and giving the aircraft a very agile control behavior even during take-off and landing. This thrust vector control makes it possible (assuming sufficient propulsion system power) to land the aircraft provided with an aerostatic lift body with essentially pinpoint accuracy. This affords the advantage that the aircraft can descend directly on relatively small outside landing areas and can thus, for example, pick up a load in a factory yard and deposit it again directly with the recipient.
The inclination of the rotor plane takes place due to aerodynamic forces acting on the propeller blades, as the result of individual adjustment of the respective angle of incidence of the individual propeller blades. In this case, only the thrust vector for propulsion, lift and control is introduced as a force into the fuselage. Reaction moments occurring during a rapid build-up of the thrust vector, for example gyroscopic moments of an associated engine or of the propeller, are supported on the surrounding air and not on the structure of the aircraft. In this way, the propeller plane is also tiltable relative to the output shaft very quickly in any direction over a large angular sector, without reaction forces which in this case originate from gyroscopic moments having to be transmitted to the fuselage. The angle of inclination of the propeller rotation plane relative to the output shaft of the associated drive, said output shaft acting on the propeller shaft, may amount to between xc2x120xc2x0 and xc2x150xc2x0, preferably between xc2x125xc2x0 and xc2x135xc2x0 and, for further preference, xc2x130xc2x0. Since the thrust vector control of the aircraft according to the invention works both in the propulsion position and in the lift position of the propulsion units, complicated restraint of the aircraft on the ground is not necessary for short landing stops, with the propulsion systems still running, since the rapidly reacting thrust vector control allows stabilization of the position of the aircraft on the landing area, even in the event of crosswind or gusts of wind. As a result, the aircraft according to the invention becomes independent of landing platforms or other landing equipment provided on the ground, such as, for example, anchors for securing restraining lines for the aircraft. It may nevertheless be necessary, particularly when the aircraft stops for longer periods during which the engines are switched off, to restrain the aircraft on the ground in the known way. This may be carried out by anchoring a landing foot preferably integrated on the underside of the aircraft, or by means of a rope winch system which is integrated into the fuselage and which can be activated preferably centrally. The rapid-reaction thrust vector control of the aircraft according to the invention also makes it possible to pick up a load, and deposit it with pinpoint accuracy, from the hovering state of the aircraft, without the latter itself having to land.
In a particularly preferred embodiment of the aircraft according to the invention, the fuselage is designed at the same time as an aerodynamic lift body. As a result, during cruising, the fuselage can also generate an aerodynamic lift force in addition to the aerostatic lift.
In the aircraft according to the invention, actively actuable aerodynamic control devices dependent on the dynamic pressure in flight may be dispensed with, so that there is no need to provide any horizontal elevator or rudder units which would increase the crosswind sensitivity and gust sensitivity of the aircraft, even though they do not essentially cooperate in controlling the aircraft in flight at low speed in the take-off phase and the landing phase, which is why this control is performed solely by the thrust vector control. The aircraft may thereby be designed with a consistently simple shape.
If the fuselage has an essentially circular plan, then, on the one hand, assuming the same length, because of the larger volume a substantially increased lift is achieved, as compared with the cigar-like shape of an airship of conventional type, thus leading indirectly to a higher payload, and, on the other hand, assuming the same volume, the wetted surface of the outer skin is reduced, thus leading to a weight reduction and diminishing the frictional resistance. Moreover, crosswind sensitivity is thereby substantially reduced.
It is advantageous, at the same time, if the fuselage has an essentially elliptic cross section, with the result that the flow resistance in horizontal flight is markedly reduced.
If the fuselage cross section is designed with an asymmetric, essentially elliptic shape, the upper part forming an upper shell which is curved to a greater extent than the flatter lower part forming a lower shell, then, during cruising, the fuselage also contributes, in the horizontal position, not only to the aerostatic lift, but additionally to an aerodynamic lift which replaces the rotor lift used during take-off and landing. Moreover, the different curvature leads to a more balanced structural load in the upper shell and the lower shell of the asymmetric discus-like fuselage.
Preferably, the fuselage has, in the equatorial region, at least one rim-like reinforcing ring which forms a horizontal stiffening of the fuselage, in that said reinforcing ring absorbs the radial forces of the upper shell and of the lower shell and additionally, for example, supports dynamic pressure forces impinging on the fuselage on the leading edge side.
At the same time, it is advantageous, in particular, if the reinforcing ring has, in cross section, a part-elliptic shape on its outer circumference. The reinforcing ring is thereby adapted, in cross section, to the shape of the fuselage cross section in the equatorial region.
The embodiment is also advantageous in which the reinforcing ring has a composite fiber material, preferably of the sandwich type. This achieves, along with high strength and low weight, a desired elasticity which allows the reinforcing ring to experience deformation within predetermined limits. By virtue of this elastic deformation, for example, forces and moments introduced into the reinforcing ring from a propulsion support structure can be taken over and transferred by the fuselage envelope structure which, on account of the largest possible lever arms and the natural dimensional rigidity of the pressurized elliptic discus-like fuselage, is particularly suitable for this purpose. The main function of the rim-like reinforcing ring is to absorb the radial forces from the envelope structure for the purpose of the horizontal stiffening of the discus-like fuselage.
It is advantageous for the resultant compression loads to be absorbed by two compression-resistant supporting profiles with high specific compressive strength, which are integrated into the reinforcing ring and to which preferably also the support structure of the upper shell and of the lower shell of the fuselage is anchored.
In order to limit the desired radial elasticity, the reinforcing ring may have at least one supporting skeleton which i s advantageously designed as a framework which, in cross section, is preferably of essentially triangular design, two of the corners being formed by the supporting profiles integrated into the reinforcing ring, and the vertex of the triangle pointing toward the inside of the fuselage.
It is advantageous if the supporting skeleton is integrated at least partially into the reinforcing ring.
If, according to a further advantageous embodiment of the invention, in each case two propulsion units are jointly mounted in a supporting structure preferably attached nonrigidly to the fuselage, then bending moments, which result, during the take-off and landing phases, from the lifting thrust of the individual propulsion system and the projecting attachment of the propeller axis, can be guided directly from one propulsion unit to the other, without these forces having to be transmitted through the entire fuselage structure. The two drive units are thus supported relative to one another via their supporting structure.
The respective propulsion units and/or their supporting structures together with the propulsion units assigned to them are preferably coupled nonrigidly to one another via thrust struts, to form a propulsion support frame which is twistable and distortable as a result of its nonrigid connections.
Preferably, the propulsion units and/or their supporting structures are attached nonrigidly to the rim-like reinforcing ring. The propulsion units are thereby integrated into the propulsion support frame which is attached nonrigidly to the rim-like reinforcing ring. This embodiment ensures that the forces emanating from the individual propulsion units are largely transferred via the propulsion support frame and therefore the reinforcing ring and the fuselage envelope structure are relieved of the transfer of these forces. In addition, vibrational uncoupling is assisted thereby.
If the front propulsion units and the rear propulsion units are in each case located at a different distance from the longitudinal center plane, this ensures that the rear propulsion units do not lie in the vortex trail of the front propulsion units.
Additionally or alternatively to this, the front and rear propulsion units may also be arranged at different heights on the aircraft, in order to achieve the same or a more improved effect in this respect.
In a particularly preferred design of the invention, the aircraft is provided with four propulsion units which, for further preference, are in each case provided in pairs on a supporting structure. Advantageously, in each case, a propulsion unit is provided in the region of one corner of an imaginary quadrangle (or another polygon, depending on the number of propulsion units) which surrounds or partially penetrates the circular plan of the aircraft.
Preferably, two engines capable of being operated in parallel with one another are provided in each propulsion unit.
Redundancy is thereby achieved within each individual propulsion unit, and, even if an engine of a propulsion unit fails, this redundancy also allows the propulsion unit as a whole to operate reliably, with only a slight overall loss of thrust. The operating safety of the aircraft is thereby increased, since the risk of a complete failure of an entire propulsion unit is greatly reduced because of the duplication of the engines. The arrangement of four twin-engine propulsion units of this type gives full propulsion redundancy, even if an engine were to fail during take-off with a maximum take-off mass, that is to say in vertical flight. If a complete propulsion unit fails during take-off, only two propulsion units located diagonally opposite one another provide lift and, in the case of a maximum take-off mass, make it possible to maintain an only insignificant descent speed, the third operational propulsion unit being used for stabilizing the aircraft about the roll axis and about the pitch axis. If, in the case of such a loss of propulsion, a sufficient flight altitude has already been reached, a transition to cruising can be carried out. During cruising, too, the aircraft provided with four propulsion units in the way claimed remains fully airworthy and maneuverable if a propulsion unit fails completely, since, in this case, it is ensured that there is still one propulsion unit operational on each side of the aircraft with respect to the longitudinal center plane, the third operational propulsion unit also being used for regulating the flight attitude.
A cargo compartment for the transport of cargo is preferably designed in the lower region of the fuselage, below which cargo compartment a landing foot preferably of platform-like design can be extended.
In an advantageous embodiment, the cargo compartment is provided with at least one ramp, and preferably two ramps can be provided on two sides facing away from one another. The provision of one ramp makes it easier for the aircraft to be loaded and unloaded and the provision of two ramps located on sides facing away from one another allows more rapid loading and unloading in the so-called RORO mode (roll-on/roll-off).
If a pneumatically extendable bellows-like annular bead is provided, directed downward, as a landing foot below the cargo compartment in the region of the circumference of the latter, then, on the one hand, landing impact can be cushioned by the landing foot formed by this annular bead and, on the other hand, because of the low specific surface pressure, landing on unconsolidated ground may also take place. In order to set a specific height, the pneumatically extendable annular bead has integrated height limitation.
In another embodiment of the aircraft according to the invention, a passenger cabin, preferably having a two-story design in places, is provided in the front part of the equatorial region of the fuselage, so that the aircraft can be used as a means of passenger transport.
Preferably, in this case, the passenger cabin is suspended in the rim-like reinforcing ring and preferably also in the front supporting structure of the propulsion support frame.
In a preferred embodiment, a baggage and freight compartment is provided in the rear part of the equatorial region of the fuselage. This arrangement of the baggage and freight compartment in the rear part of the aircraft ensures, together with the passenger cabin provided in the front part of the aircraft, that the basic trim of the aircraft is as balanced as possible.
Preferably, the baggage and freight compartment is suspended in the rim-like reinforcing ring and preferably also in the rear supporting structure of the propulsion support frame.
In an advantageous development, there is provided in the lower shell a central body which is integrated into the latter and on the underside of which preferably a bellows-like pneumatically extendable annular bead is designed as a landing foot.
If the central body is suspended in the envelope structure of the fuselage, said envelope structure being formed by the upper shell and by the lower shell, in such a way that, in the event of a hard landing, said central body can spring upward and thus makes it possible to cushion the passenger cabin, baggage and freight compartment and propulsion support frame, landing impacts are kept away from the passenger cabin and consequently from the passengers and also from the baggage and freight compartment and the propulsion support frame.
Preferably, the central body is provided with at least one ramp for access from outside.
If the passenger cabin and the baggage and freight compartment are connected to the central body via encased transport links, then connecting passages screened off from the surrounding fuselage interior are produced between the central body and the passenger cabin as well as the baggage and freight compartment. If the connection between the transport links and the central body is of nonrigid design, the springing of the central body becomes possible. In this case, at least two, preferably three transport links may be provided.
In a further preferred embodiment, the fuselage has a support structure and a fuselage envelope, the fuselage envelope being heatable, at least in portions, in the region of the upper shell. This heating, in particular on that side of the upper shell which faces the inside of the fuselage, causes the fuselage envelope to be deiced, with the result that the operating safety of the aircraft in use during bad weather is appreciably increased.
Advantageously, the heatable portions of the fuselage envelope may be of double-walled design and have flowing through them, as required, warm air or another gas which is warmer than the fuselage surroundings. For this purpose, either the waste heat of the engines may be used or additional independent heating devices may be provided. It is also advantageous if the pressure within the fuselage envelope is capable of being modulated. This embodiment assists effective deicing of the fuselage envelope and is consequently conducive to the aircraft operating safely in bad weather.
In another particularly preferred embodiment of the aircraft according to the invention, central, preferably digital control is provided for the individual or collective control of the angles of incidence of the propeller blades of all the propulsion units for preferably exclusive attitude control and for flight control in the vertical take-off and landing mode, in the horizontal cruising mode and in the transitional mode between these two operating conditions. This central control ensures a stable flight behavior which is guaranteed at all the operating conditions of the aircraft and thus relieves the pilots of this task.
Additional manual flight control may be provided as redundancy for this central control, said manual flight control enabling the pilot to stabilize the flight behavior of the aircraft in the event of a failure of central control.
The thrust vector control in the aircraft according to the invention is provided by a propulsion unit having at least one propeller, in which the propeller rotation plane is designed to be inclinable relative to the drive output shaft acting on the propeller shaft, uniformity in the rotational movement of the propeller being achieved in that the propeller shaft and the drive output shaft acting on the propeller shaft are connected to one another in an articulated manner, preferably via a double cardan joint or a synchronous joint. This inventive embodiment of a propulsion unit additionally achieves, independently of the tilting state of the propeller rotation plane between the lift position and the propulsion position, all-round effective inclinability of the propeller rotation plane in the form of an imaginary disk which allows a rapid and immediately effective change in the thrust vector. This special embodiment of a propulsion unit, with a propeller rotation plane inclinable all-round relative to the drive output shaft, can be used not only on the aircraft described in this application, but for aircraft in general or, for example, also for vessels, when the effective direction of a thrust vector emanating from a rotating propulsion unit is to be changed quickly.
For practical use, preferably in aircraft, in addition to the primary suitability for use in the highly agile thrust vector control of propeller propulsion units, there are the following advantages:
The propeller plane can in each case be oriented perpendicularly to the air flow direction, irrespective of the attitude or aircraft position.
In the case of the approach of a crosswind, a compensating trimming thrust component can be built up.
Even when the air flow direction is oblique to the propeller plane, no bending moment acts on the propeller shaft, since the build-up of a tilting moment of the propeller plane is eliminated by means of the cyclically individual blade setting.
It is advantageous, at the same time, if the hub of the propeller is cardanically mounted via a cardan ring, with the result that the propeller rotation plane inclinability making thrust vector control possible is achieved.
In a preferred development of this propulsion unit, the propeller blades are arranged on an associated propeller hub without flapping hinges and without lag hinges or other elastic parts acting in an equivalent way to these.
In this case, the angle of incidence of the individual propeller blades is adjustable, preferably by means of a swashplate, collectively as well as individually variably relative to the inclination of the propeller rotation plane. This arrangement of the propeller blades and the control of their angles of incidence via a swashplate bring about an inclination of the propeller rotation plane which directly follows the change in the angle of incidence of the propeller blades (pitch change), with the result that the desired thrust vector change for controlling the aircraft is achieved.
In this case, the cyclic change in the blade angles of incidence takes place via the swashplate, as in the case of the helicopter rotor. In contrast to this, however, the blade roots of the propeller blades do not have to continue to be adjusted cyclically after the conclusion of the dynamic operation of tilting the propeller plane into a plane perpendicular to the air flow direction, since, in contrast to the rotor hub fixed to the helicopter, the propeller hub has been tilted jointly with the propeller rotation plane relative to the output shaft. In this embodiment of the subject of the invention, the swashplate, the propeller hub and the propeller blades rotate in planes parallel to one another again after the tilting operation.
Admittedly, in principle, the angle of inclination of the propeller rotation plane relative to the output shaft of the associated drive, said output shaft acting on the propeller shaft, may amount to between xc2x120xc2x0 and xc2x150xc2x0, preferably between xc2x125xc2x0 and xc2x135xc2x0 and, for further preference, xc2x130xc2x0, as has already been stated. However, if the angle of inclination of the propeller rotation plane relative to the output shaft of the associated drive, said output shaft acting on the propeller shaft, amounts up to more than xc2x145xc2x0, then, with the propulsion unit being suitably attached obliquely to a craft, both a lift position and a propulsion position can be set solely by inclining the propeller rotation plane.
Preferably, however, a tilting mechanism for mounting the propulsion unit on a craft is provided, said tilting mechanism allowing the propulsion unit to tilt about a tilting axis between a lift position, in which the output shaft is oriented essentially vertically, and a propulsion position, in which the output shaft is oriented essentially horizontally. This tilting of the propeller rotation plane out of the horizontal position (lift position) into the vertical position (propulsion position), and vice versa, during the transitional phases, that is to say, for example, in the case of the aircraft, the respective transitional phases between vertical flight and horizontal flight, is likewise induced by the individual control of the angles of incidence of the propeller blades via fluid-dynamic forces and causes the output axis of the propulsion system to tilt about the tilting axis, for example an axis parallel to the transverse axis of the aircraft.
Preferably, a tracking device is provided, which follows a tilting movement of the propulsion unit, in particular the propeller rotation plane, occurring due to fluid-dynamic forces acting on the propeller and to resultant gyroscopic forces, and which assists this tilting movement, preferably without any reaction force. The tracking device, in this case, follows, with a markedly lowered adjusting speed (approximately the factor 5), the tilting movement of the propeller rotation plane which commences due to the fluid-dynamic forces acting on the propeller (these are aerodynamic forces in use on an aircraft) and due to the resultant gyroscopic forces.
In an alternative embodiment, the propeller hub is mounted in a uniaxial inclination joint, the inclination axis of which runs perpendicular to the tilting axis of the propulsion unit, so that the inclinability of the propeller hub about the inclination axis, together with the tiltability of the propulsion unit about the tilting axis of the propulsion unit, allows the propeller rotation plane to be inclined in all directions, the adjusting speed of the tracking device for the tilting movement about the tilting axis of the propulsion unit essentially corresponding to the adjusting speed of the fluid-dynamically induced inclining movement of the propeller rotation plane, in order to achieve a tilting movement essentially free of reaction force. The cardanic mounting, specified further above, of the propeller hub may be dispensed with in this design.
In a preferred development, there is integrated into the propeller hub a reduction gear which is preferably designed in the manner of a planetary gear and is acted upon rotationally by the output shaft of the drive, preferably via the double cardan joint or the synchronous joint, and which transmits the rotational speed of the output shaft, reduced, to the propeller hub. The double cardan joint or the synchronous joint for the propeller drive is thereby relieved of very high moments which may occur, in particular, during operation with a propeller of large diameter.
This embodiment of the propulsion unit according to the invention, because of its propeller rotation plane inclination caused by the variable-pitch setting of the angle of incidence of the rotor blades, ensures that the propeller rotation plane is deflected out of its current position without any reaction force and, consequently, that the thrust vector is changed without any reaction force. In this embodiment, therefore, no gyroscopic moments have to be supported on the fuselage, so that complicated and heavy supporting structures and corresponding reinforcements in the fuselage may also be dispensed with, even when rapid thrust vector changes are required for agile flight control and attitude control.
The invention relates, furthermore, to a method for controlling an aircraft having a propeller propulsion system, the angle of incidence of the individual propeller blades of each propeller being cyclically set individually, and, thereupon, the propeller rotation plane being inclined without any reaction force, said inclination being induced by aerodynamic forces and by gyroscopic forces resulting from these. This method allows the rapid-reaction control of a propeller aircraft which as a result, particularly in the low flying speed range, allows more rapid changes in direction than with the conventional aerodynamic control via elevators, rudders and ailerons.